Fluid propulsion devices achieve thrust by imparting momentum to a fluid called the propellant. An air-breathing engine, as the name implies, uses the atmosphere for most of its propellant. The gas turbine produces high-temperature gas which may be used either to generate power for a propeller, fan, generator or other mechanical apparatus or to develop thrust directly by expansion and acceleration of the hot gas in a nozzle. In any case, an air breathing engine continuously draws air from the atmosphere, compresses it, adds energy in the form of heat, and then expands it in order to convert the added energy to shaft work or jet kinetic energy. Thus, in addition to acting as propellant, the air acts as the working fluid in a thermodynamic process in which a fraction of the energy is made available for propulsive purposes or work.
Typically gas turbine engines include at least two air streams. All air utilized by the engine initially passes through a fan, and then it is split into the two air streams. The inner air stream is referred to as core air and passes into the compressor portion of the engine, where it is compressed. This air then is fed to the combustor portion of the engine where it is mixed with fuel and the fuel is combusted. The combustion gases are then expanded through the turbine portion of the engine, which extracts energy from the hot combustion gases, the extracted energy being used to run the compressor, the fan and other accessory systems. The remaining hot gases then flow into the exhaust portion of the engine, which may be used to produce thrust for forward motion to the aircraft.
The outer air flow stream bypasses the engine core and is pressurized by the fan. Typically, no other work is done on the outer air flow stream which continues axially down the engine but outside the core. The bypass air flow stream also can be used to accomplish aircraft cooling by the introduction of heat exchangers in the fan stream. Downstream of the turbine, the outer air flow stream is used to cool engine hardware in the exhaust system. When additional thrust is required (demanded), some of the fans bypass air flow stream may be redirected to the augmenter (afterburner) where it is mixed with core flow and fuel to provide the additional thrust to move the aircraft.
Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views, FIG. 1 shows a general orientation of a turbofan engine in a cut away view. In the turbofan engine shown, the flow of the air is generally axial. The engine direction along the axis is generally defined using the terms “upstream” and “downstream” generally which refer to a position in a jet engine in relation to the ambient air inlet and the engine exhaust at the back of the engine. For example, the inlet fan is upstream of the combustion chamber. Likewise, the terms “fore” and “aft” generally refer to a position in relation to the ambient air inlet and the engine exhaust nozzle. Additionally, outward/outboard and inward/inboard refer to the radial direction. For example the bypass duct is outboard the core duct. The ducts are generally circular and co-axial with each other.
As ambient inlet airflow 12 enters inlet fan duct 14 of turbofan engine 10, through the guide vanes 15, passes by fan spinner 16 and through fan rotor (fan blade) 42. The airflow 12 is split into primary (core) flow stream 28 and bypass flow stream 30 by upstream splitter 24 and downstream splitter 25. The bypass flow stream 30 along with the core/primary flow stream 28 is shown, the bypass stream 30 being outboard of the core stream 28. The inward portion of the bypass stream 30 and the outward portion of the core streams are partially defined by the splitters upstream of the compressor 26. The fan 42 has a plurality of fan blades.
As shown in FIG. 1 the fan blade 42 shown is rotating about the engine axis into the page, therefor the low pressure side of the blade 42 is shown, the high pressure side being on the opposite side. The primary flow stream 28 flows through compressor 26 that compresses the air to a higher pressure. The compressed air typically passes through an outlet guide vane to straighten the airflow and eliminate swirling motion or turbulence, a diffuser where air spreads out, and a compressor manifold to distribute the air in a smooth flow. The core flow stream 28 is then mixed with fuel in combustion chamber 35 and the mixture is ignited and burned. The resultant combustion products flow through turbines 38 that extract energy from the combustion gases to turn fan rotor 42, compressor 26 and any shaft work by way of turbine shaft 40. The gases, passing exhaust cone, expand through an exhaust nozzle 43 to produce thrust. Primary flow stream 28 leaves the engine at a higher velocity than when it entered. Bypass flow stream 30 flows through fan rotor 42, flows by bypass duct outer wall 27, an annular duct concentric with the core engine flows through fan discharge outlet and is expanded through an exhaust nozzle to produce additional thrust. Turbofan engine 10 has a generally longitudinally extending centerline represented by engine axis 46.
The turbine inlet temperature is very high and the turbine requires materials with special heat resistant and strength characteristics. The turbine blades are held by the turbine disc and the turbine discs are typically do not possess all of the characteristics of the blades and should be protected from the hot gases passing through the turbine blades. Additionally the flow of cooling gases is advantageously controlled throughout the engine. Full face cover plates are used to isolate the turbine discs from the hot gasses, retain the turbine blades and establish flow paths for cooling gases. These cover plates are typically provided for each of the turbine discs and in particular the high pressure turbine stages with cooled blades, where the gases are the hottest.
Full face cover plates are typically assembled with the turbine disc using a retaining ring, bayonet features or a bolted flange. However some configurations due to space or access cannot accommodate a flange or bayonet features, and similarly due to the size of the wheel may not be sufficient to allow for a retaining ring.
FIG. 2 illustrates a cutaway of a turbine disc 200. The forward face of the disc 200 is covered by full face cover plate 202, full face cover plate 202 being held in place by retaining ring 204. A retaining ring may also be used to secure a cover on the aft face of the disc 200. Installing retaining rings, such as ring 204, requires special tooling, which adds cost and complexity, and sufficient assembly access which may be limited in applications of turbines with small diameters.
FIG. 3 illustrates a cutaway of a turbine disc 300. The forward face of disc 300 is covered by full face cover plate 302, full face cover plate 302 being held in place by bolted flange 304. Using a bolted flange 304 adds additional components and weight to the assembled turbine system. This increase in weight tends to cause a bolted flange to be heavier than using a retaining ring. A bolted flange also requires sufficient space for wrench access during assembly which may be limited or not available depending on the applications. Additionally, a bolted flange reduces engine efficiency by creating windage losses. These losses can be eliminated by using a shield such as shield 306. However, this additional component further increases costs, complexity of assembly and even more weight.
FIG. 4 illustrates a cutaway of a turbine disc 400. The forward face of the disc 400 is covered by partial face cover plate 402. This partial face cover plate 402 is held in place using bayonet features 404. These bayonet features 404 require additional machining of the cover plate 402 and disc 400 and are not axisymmetric which can lead to high local stresses. Additionally, special tooling is needed to install cover plate 402 using bayonet features 404 adding to the cost and complexity of turbine assembly.
As disclosed in some embodiments herein the current subject matter addresses these deficiencies by utilizing a spanner nut and buttress threads on the turbine wheel drive arm to axially restrain the full face cover plate against the turbine disc rim face. The current subject matter requires no special assembly tooling (as is required for retaining ring and bayonet features assemblies) is lighter weight and has fewer parts than a bolted flange, and offers a simplified assembly process, simplified machining, variable and repeatable load control and no high stress 3D features.
The disclosed subject matter in accordance with some embodiments also addresses these deficiencies between the turbine disc of a multi-stage turbine by utilizing the tie bolt and spanner nut assembly to clamp the cover plated between an axial stop on the first stage disc and the rim of the second stage disc. This approach also requires no special assembly tooling (as is required for retaining ring and bayonet features assemblies) is lighter weight and has fewer parts than a bolted flange, and offers a simplified assembly process, simplified machining, variable and repeatable load control and no high stress 3D features.
In accordance with some embodiments of the present disclosure, a gas turbine engine is presented. The engine has at least one turbine stage and each stage has a turbine disc which is concentric with the axis of the engine. The turbine disc has a turbine arm positioned axially from the disc and rigidly attached thereto. The drive arm is also concentric with the turbine axis and has an outer surface. The turbine stage also has a cover plate concentric with the turbine axis and covering at least a portion of the turbine disc. An arrangement of turbine blades are positioned around the periphery of the disc and are retained on the turbine disc at least partially by the cover plate. The system includes a spanner nut and threads on a portion of the turbine drive arm. The spanner nut is threaded onto the turbine arm which applies an axial force on the cover plate, thereby retaining the cover plate.
In accordance with some embodiments of the present disclosure, a turbine assembly is provided. The assembly includes a turbine disc having a center axis and a radially outer rim, a turbine disc arm coaxial with the turbine disc and extending axially from the disc and having a threaded portion, a cover plate having a first end and a second end which is coaxial with the turbine disc and a spanner nut coaxial with the turbine disc. The spanner nut is threaded onto the turbine disc arm and contacts the first end of the turbine cover plate and holds the second end in contact with the outer rim.
In accordance with some embodiments of the present disclosure, a method of retaining a cover plate is provided. A gas turbine engine comprises an axis, turbine disc and a cover plate. The turbine disc is coaxial with the axis and includes an arm extending axially from the turbine disc. The turbine disc also includes a rim around a periphery of the disc. The cover plate defines a first and second end concentric openings, each opening with a radius greater than the radius of the arm. The arm of the turbine disc is threaded to produce a threaded arm. The cover plate is moved axially such that the arm is within the openings. The second end of the cover plate is positioned proximate to the rim of the turbine disc. A rotation is applied to the spanner nut, thereby threading the spanner nut onto the threaded arm. This rotation is coaxial with the axis. The spanner nut contacts the cover plate and applies a force via the contact to retain the second end against the rim.
These and many other advantages of the present subject matter will be readily apparent to one skilled in the art to which the invention pertains from a perusal of the claims, the appended drawings, and the following detailed description of preferred embodiments.